Aircraft power plant

ABSTRACT

Aircraft power plants and associated methods are provided. A method for driving a load on an aircraft includes: transferring motive power from an internal combustion (IC) engine to the load; discharging a flow of first exhaust gas from the IC engine when transferring motive power from the IC engine to the load; receiving the flow of first exhaust gas from the IC engine into a combustor; mixing fuel with the first exhaust gas in the combustor and igniting the fuel to generate a flow of second exhaust gas; receiving the flow of second exhaust gas at a turbine and driving the turbine with the flow of second exhaust gas from the combustor; and transferring motive power from the turbine to the load.

TECHNICAL FIELD

The disclosure relates generally to aircraft, and more particularly toaircraft power plants.

BACKGROUND

It is desirable for aircraft engines to operate in an energy efficientmanner to promote reduced fuel consumption and operating costs. Inaircraft applications, the power output capacity of an engine relativeto the weight of the engine is an important factor that can affect theoverall efficiency of the aircraft since the weight of the engine mustbe carried by the aircraft during flight. Aircraft engines can berequired to have a maximum power output rating that can be producedduring short-term (i.e., momentary) peak power operation in situationssuch as during take-off or during emergency situations. Even though, thelong-term continuous operation of the aircraft engines during a cruisephase of flight, for example, can be well below such maximum poweroutput rating, the maximum power output rating can necessitate increasedsize and weight of aircraft engines. Improvement is desirable.

SUMMARY

In one aspect, the disclosure describes an aircraft power plantcomprising:

-   -   an internal combustion (IC) engine using intermittent combustion        during operation, the IC engine being drivingly connected to a        load to transfer motive power from the IC engine to the load,        the IC engine having an exhaust outlet for discharging a flow of        first exhaust gas;    -   a combustor including: a combustor inlet in fluid communication        with the exhaust outlet of the IC engine to receive the flow of        first exhaust gas from the IC engine into the combustor; a fuel        injector for injecting fuel into the combustor in which the fuel        is mixed with the first exhaust gas and ignited to generate a        flow of second exhaust gas; and a combustor outlet for        discharging the flow of second exhaust gas; and    -   a turbocharger associated with the IC engine, the turbocharger        including: a turbine in fluid communication with the combustor        outlet and driven by the flow of second exhaust gas, the turbine        being drivingly connected to the load to transfer motive power        from the turbine to the load; and a compressor drivingly        connected to the turbine and driven by the turbine to compress        combustion air for the IC engine.

In another aspect, the disclosure describes an aircraft engine systemcomprising:

-   -   a Wankel engine drivingly connected to a load to transfer motive        power from the Wankel engine to the load, the Wankel engine        having an exhaust outlet for discharging a flow of exhaust gas        from the Wankel engine;    -   a turbine in fluid communication with the exhaust outlet of the        Wankel engine and driven by the flow of exhaust gas, the turbine        being drivingly connected to the load to transfer motive power        from the turbine to the load; and    -   an inter-burner operatively disposed between the Wankel engine        and the turbine to add energy to the flow of exhaust gas between        the Wankel engine and the turbine.

In a further aspect, the disclosure describes a method of driving a loadonboard an aircraft. The method comprises:

-   -   transferring motive power from an internal combustion (IC)        engine to the load, the IC engine using intermittent combustion        during operation;    -   discharging a flow of first exhaust gas from the IC engine when        transferring motive power from the IC engine to the load;    -   receiving the flow of first exhaust gas from the IC engine into        a combustor;    -   mixing fuel with the first exhaust gas in the combustor and        igniting the fuel to generate a flow of second exhaust gas;    -   receiving the flow of second exhaust gas at a turbine and        driving the turbine with the flow of second exhaust gas from the        combustor; and    -   transferring motive power from the turbine to the load.

Further details of these and other aspects of the subject matter of thisapplication will be apparent from the detailed description includedbelow and the drawings.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying drawings, in which:

FIG. 1 schematically shows an aircraft including a twin-engine powerplant as described herein;

FIG. 2 shows an exemplary schematic representation of an engine systemof the aircraft power plant of FIG. 1 ;

FIG. 3 is a flowchart of an exemplary method of driving a load onboardan aircraft;

FIG. 4 is a graphical representation of different operating regimes ofthe engine system of FIG. 2 together with exemplary sizing requirementsfor elements of the engine system; and

FIG. 5 is another graphical representation of different operatingregimes of the engine system of FIG. 2 together with other exemplarysizing requirements for elements of the engine system.

DETAILED DESCRIPTION

The present disclosure describes aircraft power plants, engine systems,and associated methods. In various embodiments, aircraft power plantsdescribed herein may be configured to provide momentary supplementalpower output without necessitating significant size or weight increaseof the aircraft power plants. For example, the aircraft power plantsdescribed herein may include an internal combustion (IC) engine with oneor more associated systems such as a liquid cooling system, an oilcooling system, and bearings that are sized according to a power outputcapacity of the IC engine. In some embodiments, the momentarysupplemental power output capacity of the power plant may be providedvia a combustor (i.e., inter-burner) in which fuel may be mixed withexhaust gas from the IC engine, and the mixture may be ignited togenerate supplemental power that may be extracted via a turbine and usedwhen needed. In some embodiments, the combustor may provide momentarysupplemental power output capacity for the power plant as a wholewithout increasing the power output capacity of the IC engine, the sizeand weight the (e.g., liquid cooling) systems of the IC engine, andbearings of the IC engine. Accordingly, the IC engine and its associatedsystems may be sized and designed for efficient operation duringlong-term continuous operation of the aircraft power plant during acruise phase of flight for example, and the supplemental power outputcapacity may be provided by selectively activating the combustor whenneeded.

Aspects of various embodiments are described through reference to thedrawings.

The term “connected” may include both direct connection (where twoelements contact each other) and indirect connection (where at least oneadditional element is located between the two elements).

The term “substantially” as used herein may be applied to modify anyquantitative representation which could permissibly vary withoutresulting in a change in the basic function to which it is related.

FIG. 1 schematically shows aircraft 10 including aircraft power plant 12(referred hereinafter as “power plant 12”) as described herein. Invarious embodiments, aircraft 10 may be a manned aircraft or an unmannedaircraft (e.g., drone). For example, aircraft 10 may be a corporate,private, fixed-wing, rotary-wing (e.g., helicopter) or commercialpassenger aircraft. In some embodiments, power plant 12 may be part of apropulsion system of aircraft 10. Power plant 12 may be drivinglyconnected to one or more loads 14 (referred hereinafter in the singular)via output shaft 16. In some embodiments, load 14 may include apropeller configured to generate thrust and/or lift for aircraft 10. Insome embodiments, load 14 may include a rotary wing (e.g., main rotor)of a helicopter or other type of rotorcraft. In some embodiments, powerplant 12 may provide energy for propelling aircraft 10, and optionallyalso provide energy to drive loads for functions other than propulsionsuch as driving a compressor supplying compressed air to pneumaticloads, driving an electric generator supplying electric energy toelectric loads, driving an oil pump supplying oil to lubrication loads,and/or driving a hydraulic pump supplying pressurized hydraulic fluid tohydraulic actuators. Alternatively, power plant 12 may be an auxiliarypower unit (APU) of aircraft 10 and may provide energy exclusively forfunctions other than propulsion of aircraft 10.

Aircraft power plant 12 may include a single engine system 18A, or aplurality of engine systems 18A, 18B. For example, power plant 12 may bea multi-engine system (e.g., twin-pack power plant) including firstengine system 18A and second engine system 18B. Control of first andsecond engine systems 18A, 18B may be effected by one or morecontroller(s) 20, which may be one or more full authority digital enginecontrol(s) (“FADEC(s)”), electronic engine controller(s) (EEC(s)), orthe like, that are programmed to manage the operation of engine systems18A, 18B during various phases of flight of aircraft 10 to reduce anoverall fuel burn for example. Controller(s) 20 may include a firstcontroller for controlling first engine system 18A and a secondcontroller for controlling second engine system 18B. In someembodiments, a single controller 20 may be used for controlling firstengine system 18A and second engine 18B.

Power plant 12 may be operable so that engine systems 18A, 18B may beoperated symmetrically or asymmetrically, with one engine systemoperating in a high-power “active” mode and the other engine systemoperating in a lower-power (which could be no power, in some cases)“standby” mode. Doing so may provide fuel saving opportunities toaircraft 10. However there may be other reasons why engine systems 18A,18B may be operated asymmetrically. Such asymmetric operation of enginesystems 18A, 18B may be engaged for a cruise phase of flight(continuous, steady-state flight which is typically at a given commandedconstant aircraft cruising speed and altitude) for example. Power plant12 may be used in aircraft applications but may also be suitable formarine, industrial and/or other ground operations.

In use, first engine system 18A may operate in the active mode whilesecond engine system 18B may operate in the standby mode. During thisasymmetric operation, if the helicopter needs a power increase (expectedor otherwise), second engine system 18B may be required to provide morepower relative to the low power conditions of the standby mode, andpossibly return immediately to a high- or full-power condition. This mayoccur, for example, during take-off or in an emergency condition wherethe “active” first engine system 18A loses power. Even absent anemergency, it may be desirable to repower the standby engine to exit theasymmetric mode in some situations.

First and second engine systems 18A, 18B may be drivingly connected toload 14 via gearbox 22, which may be of a speed-changing (e.g.,reducing) type. For example, gear box 22 may have a plurality of inputsto receive motive power (mechanical energy) from respective shafts ofengine systems 18A, 18B. Gear box 22 may be configured to direct atleast some of the combined motive power from the plurality of enginesystems 18A, 18B toward common output shaft 16 for driving load 14 at asuitable operating (e.g., rotational) speed.

FIG. 2 shows an exemplary schematic representation of first enginesystem 18A of power plant 12 shown in FIG. 1 . In some embodiments,first engine system 18A and second engine system 18B of power plant 12may be of substantially identical construction, and may havesubstantially identical power output ratings. Alternatively, in someembodiments, first engine system 18A and second engine system 18B may beof different constructions and may have different power output ratings.

First engine system 18A may include internal combustion (IC) engine 24using intermittent combustion during operation. IC engine 24 may be of atype other than a gas turbine engine. For example, IC engine 24 may beof a type where the exhaust gas from IC engine 24 produces nosignificant thrust for propelling aircraft 10. Such IC engine 24 withintermittent combustion may be a reciprocating engine such as pistonengine, or a pistonless rotary (e.g., Wankel) engine where heat is addedat substantially constant volume in the thermodynamic cycle. In someembodiments, IC engine 24 may include a Wankel engine using an eccentricrotary design to convert pressure into rotating motion. In someembodiments, IC engine 24 may of a type described in U.S. Pat. No.10,107,195 (Title: COMPOUND CYCLE ENGINE), the entire contents of whichare incorporated by reference herein. In some embodiments, IC engine 24may operate on a mixture of relatively heavy fuel (e.g. diesel, kerosene(jet fuel), equivalent biofuel) and air.

IC engine 24 may drive engine output shaft 26, which may be drivinglyconnected to load 14 to transfer motive power from IC engine 24 to load14. Load 14 may include main rotor 27 of a helicopter. IC engine 24 mayinclude combustion chamber 28 including one or more fuel injectors 30A(referred hereinafter in the singular) for delivering fuel WF1 intocombustion chamber 28. Fuel WF1 may be mixed with air in combustionchamber 28 and ignited to generate a flow of first exhaust gas. The flowof first exhaust gas may be discharged from exhaust outlet 32 of ICengine 24. Exhaust outlet 32 may be an outlet of combustion chamber 28.

In some embodiments, IC engine 24 may be turbocharged by way of optionalturbocharger 34. Turbocharger 34 may include compressor 36 and turbine38 which may be drivingly interconnected via turbine shaft 40 so thatcompressor 36 may be driven by turbine 38. Compressor 36 and turbine 38may each be a single-stage device or a multiple-stage device with asingle shaft, or split on multiple independent shafts in parallel or inseries, and may be a centrifugal, axial or mixed device. Compressor 36of turbocharger 34 may receive (e.g., ambient) air and compress the airto be supplied as combustion air to IC engine 24. The ambient air may bereceived via ambient air inlet 42 and optional variable-orientationguide vanes 44A that may guide the flow of ambient air being deliveredto compressor 36 to promote a desirable operation of compressor 36 atdifferent operating conditions. In some embodiments, the compressed airfrom compressor 36 may be passed through intercooler 46. Intercooler 46may include a suitable (e.g., air to air) heat exchanger that mayfacilitate heat transfer from the compressed air to ambient air or toanother fluid to remove some of the heat added to the compressed airduring compression, and promote a desirable operation of IC engine 24.

Turbine shaft 40 of turbocharger 34 may be drivingly connected to load14 via output shaft 16, gearbox 22, and optional variable speedtransmission 48A so that motive power extracted by turbine 38 may betransferred to load 14. Variable speed transmission 48A, may permitmotive power to be transmitted from turbine shaft 40 to load 14 atvarying relative rotational speeds between turbine shaft 40 and load 14.In some embodiments, variable speed transmission 48A may be acontinuously variable transmission (CVT) that can change seamlessly andcontinuously through a range of gear ratios. Turbine shaft 40 ofturbocharger 34 may optionally be drivingly connected to one or moreauxiliary loads 14A, directly, via a clutch, via a gearbox, and/or viaanother or the same variable speed transmission 48A, so that motivepower may be transferred from turbine 38 to auxiliary load(s) 14A. Suchauxiliary load(s) 14A may include an electric generator, an oil pump ofa lubrication system, and/or a hydraulic pump of a hydraulic system forexample.

In some embodiments, IC engine 24 may be liquid cooled using suitablecoolant fluid (e.g., a solution of a suitable organic chemical such asethylene glycol, diethylene glycol, or propylene glycol in water). Forexample, heat generated by IC engine 24 may be transferred to a suitablecoolant fluid (e.g., liquid) circulating into IC engine 24, and carriedout of IC engine 24 by the coolant fluid. Heat may then be rejected bythe coolant fluid via coolant heat exchanger 50 (referred hereinafter as“coolant HX 50”), which may promote heat transfer from the coolant fluidto ambient air. The ambient air may be driven through coolant HX 50 byfan 52. Fan 52 may be configured and positioned to pull or push theambient air through coolant HX 50. In some embodiments, fan 52 may bemechanically driven by engine output shaft 26 either directly or viavariable speed transmission 48B which may be a CVT for example. However,fan 52 could instead be driven by other means such as an electric motoror a hydraulic motor for example. In some embodiments,variable-orientation guide vanes 44B may guide the flow of ambient airbeing delivered to fan 52 to promote a desirable operation of fan 52 atdifferent operating conditions.

Some heat generated by IC engine 24 may also be transferred to asuitable lubricating fluid (e.g., oil) circulating into IC engine 24,and carried out of IC engine 24 by the lubricating fluid. Heat may thenbe rejected by the lubricating fluid via oil heat exchanger 54 (referredhereinafter as “oil HX 54”), which may promote heat transfer from thelubricating fluid to ambient air. The ambient air may be driven throughoil HX 54 by fan 52 or another fan. Fan 52 may be configured andpositioned to pull or push the ambient air through oil HX 54.

First engine system 18A may include combustor 56 operatively disposedbetween IC engine 24 and turbine 38 of turbocharger 34. Combustor 56 maybe operated as an inter-burner as needed when supplemental power outputcapacity from first engine system 18A is required. When the supplementalpower output capacity from first engine system 18A is not required,combustor 56 may be inactive and the first exhaust gas discharged fromexhaust outlet 32 of IC engine 24 may be directed to turbine 38 and maydrive turbine 38. In other words, turbine 38 may extract motive powerfrom the first exhaust gas discharged from exhaust outlet 32. Duringthis mode of operation, the first exhaust gas discharged from exhaustoutlet 32 may pass through combustor 56 while combustor 56 is inactive(i.e., unlit). In some embodiments, IC engine 24 may provide a pulsatingexhaust gas flow out of exhaust outlet 32.

Combustor 56 may include combustor inlet 58 in fluid communication withexhaust outlet 32 of IC engine 24 to receive the flow of first exhaustgas from IC engine 24 into combustor 56. Combustor 56 may be disposedoutside and downstream of combustion chamber 28 of IC engine 24.Combustor 56 may be disposed outside of IC engine 24. In some operatingconditions, the air-fuel mixture that is burned in combustion chamber 28of IC engine 24 may have an air-fuel ratio that is higher thanstoichiometric (i.e., lean), and the resulting first exhaust gasdischarged from exhaust outlet 32 may contain sufficient oxygen tosustain subsequent combustion in combustor 56 in order to generate thesupplemental power output capacity of first engine system 18A. In someembodiments, first engine system 18A may include one or more valves 60that are operatively disposed to permit compressed air produced bycompressor 36 to be selectively injected into the first exhaust gasdischarged from exhaust outlet 32. Valve(s) 60 may include a three-wayvalve configured to permit a portion of the compressed air fromcompressor 36 to be delivered to intercooler 46, and another portion ofthe compressed air from compressor 36 to be injected into the firstexhaust gas downstream of exhaust outlet 32. In some situations, theadditional air may be added to the first exhaust gas if required tofacilitate subsequent combustion in combustor 56, or to help/facilitatethe (e.g., surge free) operability of compressor 36 under certainoperating conditions.

Combustor 56 may include one or more fuel injectors 30B for injectingfuel into combustor 56 where the fuel is mixed with the first exhaustgas discharged from exhaust outlet 32, and ignited for generating a flowof second exhaust gas containing a greater amount of energy. The numberand configuration of fuel injector(s) 30B may be selected to provideadequate priming and fuel delivery to combustor 56 in order to providean acceptable activation/response time in case of an emergency situationfor example. In some embodiments, the temperature of first exhaust gasentering combustor 56 may be sufficiently high so that ignition of thefuel injected into combustor 56 via injector(s) 30B may be automatic. Insome embodiments, combustor 56 may include one or more igniters tofacilitate the ignition of combustor 56.

Combustor 56 may include combustor outlet 62 for discharging the flow ofsecond exhaust gas from combustor 56. Turbine 38 may be in fluidcommunication with combustor outlet 62 and be driven by the flow ofsecond exhaust gas. In other words, turbine 38 may convert energy fromthe flow of second exhaust gas into motive power that can be used todrive compressor 36, load 14 and/or auxiliary load(s) 14A. The flow ofsecond exhaust gas flowing through turbine 38 may then be dischargedfrom first engine system 18A via exhaust duct 64 and muffler 66.

FIG. 3 is a flowchart of an exemplary method 100 of driving a load(e.g., load 14 and/or auxiliary load 14A) onboard aircraft 10. Method100 may be performed using first engine system 18A and/or using secondengine system 18B described herein or using other engine system(s).Aspects of power plant 12 and of first engine system 18A may beincorporated into method 100. Aspects of method 100 may be combined withother steps or actions disclosed herein. In various embodiments, method100 may include:

-   -   transferring motive power from IC engine 24 to load 14 (block        102);    -   discharging a flow of first exhaust gas from IC engine 24 when        transferring motive power from IC engine 24 to load 14 (block        104);    -   receiving the flow of first exhaust gas from IC engine 24 into        combustor 56 (block 106);    -   mixing fuel WF2 with the first exhaust gas in combustor 56 and        igniting the fuel WF2 to generate a flow of second exhaust gas        (block 108);    -   receiving the flow of second exhaust gas at turbine 38 and        driving turbine 38 with the flow of second exhaust gas from        combustor 56 (block 110); and    -   transferring motive power from the turbine to the load (block        112).

Method 100 may include transferring motive power from turbine 38 tocompressor 36 of turbocharger 34 to drive compressor 36. Compressor 36may compress ambient air and the compressed air may be delivered to ICengine 24 and used as compressed air by IC engine 24.

In some embodiments of method 100, load 14 may include main rotor 27 ofa helicopter.

In some embodiments of method 100, combustor 56 may be used duringmomentary peak power operation only when the supplemental power outputcapacity beyond that of IC engine 24 is required from first enginesystem 18A or from second engine system 18B. Such momentary peak poweroperation may be performed during a takeoff phase of flight of aircraft10, and/or during an emergency situation for example. Accordingly,method 100 may, after the momentary peak power operation, include:

-   -   ceasing to mix fuel WF2 with the first exhaust gas and igniting        the fuel WF2 in combustor 56 to generate a flow of second        exhaust gas;    -   receiving the flow of first exhaust gas at turbine 38 and        driving turbine 38 with the flow of first exhaust gas; and    -   transferring motive power from turbine 38 to load 14 and/or to        compressor 36 of turbocharger 34 associated with IC engine 24.

When ceasing to mix fuel WF2 with the first exhaust gas inside ofcombustor 56, fuel flow to combustor 56 via fuel injector(s) 30B may bestopped. The flow of first exhaust gas may flow through combustor 56while combustor 56 is inactive. In order to prevent coking of residualfuel WF2 inside of fuel injector(s) 30B, a suitable forward or reversefuel purging procedure may be implemented in order to clear fuelinjector(s) 30B of residual fuel WF2. For example, compressed air may bedirected into fuel injector(s) 30B in order to push the residual fuelWF2 into combustor 56 after the delivery of fuel WF2 to fuel injector(s)30B has been stopped. Alternatively, a reverse purging procedure may beimplemented to withdraw residual fuel WF2 from fuel injector(s) 30B andcollect the residual fuel WF2 into a suitable reservoir.

FIG. 4 is a graphical representation of different operating regimes offirst engine system 18A together with exemplary sizing requirements forelements of first engine system 18A. First engine system 18A may bearranged in a turboshaft setup in a twin-engine helicopter. First enginesystem 18A may be designed for a 130% maximum shaft horsepower (SHP)capability where 0% SHP corresponds to a ground idle (GI) power rating,60% SHP corresponds to a long range cruise (LRC) power rating, and 100%SHP corresponds to the required takeoff (TO) power rating. FIG. 4 alsoshows a location of a flight idle (FI) power rating disposed between GIand LRC. FIG. 4 also shows a location of a maximum cruise (MCR) powerrating disposed between LRC and TO.

FIG. 4 shows three operating regimes of first engine system 18A disposedalong a power output scale expressed in % SHP extending from 0% to 130%.First regime R1 may correspond to a relatively low power/transientoperation where the operability (e.g., no compressor surge) may be a keyobjective and specific fuel consumption (SFC) may be less important.During first regime R1, the SHP delivered to load 14 may be provided byIC engine 24 with fuel WF1 being delivered to combustion chamber 28, andwith combustor 56 being inactive where fuel WF2 is not being deliveredto combustor 56.

Second regime R2 may correspond to normal day-to-day operation where allengines of aircraft 10 are operative. Obtaining a favorable SFC may be akey objective in second regime R2. During second regime R2, the SHPdelivered to load 14 may be provided by IC engine 24 with fuel WF1 beingdelivered to combustion chamber 28, and with combustor 56 being inactivewhere fuel WF2 is not being delivered to combustor 56.

Third regime R3 may correspond to an emergency condition such as whensecond engine system 18B is inoperative (i.e., “one engine inoperative”or “OEI”) for example, and first engine system 18A must providesupplemental power output. During third regime R3, obtaining therequired SHP output from first engine system 18A may be a key objectiveand SFC may be less important. During third regime R3, the SHP deliveredto load 14 may be provided by IC engine 24 with fuel WF1 being deliveredto combustion chamber 28, and also with combustor 56 being active wherefuel WF2 is being delivered to combustor 56 as well. The supplementalpower output required from first engine system 18A may be beyond what ICengine 24 alone is capable of providing. Accordingly, the supplementalpower output may be provided by way of combustor 56.

FIG. 4 also indicates sizing requirements for various components offirst engine system 18A. As explained above the architecture of firstengine system 18A may make the supplemental power output capacityavailable without incurring significant size or weight penalty for firstengine system 18A. A maximum power output capacity of first aircraftengine system 18A may be greater than a maximum output power capacity ofIC engine 24. For example, the supplemental power capacity may beprovided separately from IC engine 24 so that the power output capacityof IC engine 24 may be left intact. Consequently, the bearing loads ofIC engine 24, and the size and weight of the (e.g., liquid-cooling, oilcooling) systems of IC engine 24 also may not need to be increased. Inother words, IC engine 24 and its associated systems may be designed andsized to function effectively and efficiently in first regime R1 andsecond regime R2 between 0% SHP and 100% SHP. First engine system 18Amay nevertheless provide the supplemental power output capacity usingcombustor 56 when required. Accordingly, as shown in FIG. 4 , severalcomponents of engine system 18A may only need to be sized for 100% SHP(i.e., TO rating) while only some components of engine system 18A mayrequired to be sized for 130% SHP.

The upsizing of muffler 66 may be optional since muffler 66 couldpotentially be bypassed during the momentary operation in third regimeR3. Variable speed transmission 48A and gearbox 22 may be designed for100% SHP but may also have the capacity to momentarily operate at 130%SHP with some design adjustments (if required).

FIG. 5 is another graphical representation of different operatingregimes of first engine system 18A together with exemplary sizingrequirements for elements of first engine system 18A. First enginesystem 18A may be arranged in a turboshaft setup in a twin-enginehelicopter. First engine system 18A may be designed for a 115% maximumSHP capability where 0% SHP corresponds to a GI power rating, 60% SHPcorresponds to a LRC power rating, and 100% SHP corresponds to therequired TO power rating. FIG. 5 also shows a location of a FI powerrating between GI and LRC. FIG. 5 also shows a location of a MCR powerrating at 85% SHP.

FIG. 5 shows three operating regimes of first engine system 18A disposedalong a power range scale expressed in % SHP extending from 0% to 115%.First regime R1 may correspond to a relatively low power/transientoperation where the operability (e.g., no compressor surge) may be a keyobjective and specific fuel consumption (SFC) may be less important.During first regime R1, the SHP delivered to load 14 may be provided byIC engine 24 with fuel WF1 being delivered to combustion chamber 28, andwith combustor 56 being inactive where fuel WF2 is not being deliveredto combustor 56.

Second regime R2 may correspond to normal day-to-day operation (otherthan TO) where all engines of aircraft 10 are operative and obtaining afavorable SFC may be a key objective. During second regime R2, the SHPdelivered to load 14 may be provided by IC engine 24 with fuel WF1 beingdelivered to combustion chamber 28, and with combustor 56 being inactivewhere fuel WF2 is not being delivered to combustor 56.

Third regime R3 may include the TO rating (100% SHP) and also anemergency condition beyond the TO rating such as when second enginesystem 18B is inoperative (i.e., “one engine inoperative” or “OEI”) forexample, and first engine system 18A must provide supplemental poweroutput. During third regime R3, obtaining the required SHP output fromfirst engine system 18A may be a key objective and SFC may be lessimportant. During third regime R3, the SHP delivered to load 14 may beprovided by IC engine 24 with fuel WF1 being delivered to combustionchamber 28, and with combustor 56 being active where fuel WF2 is beingdelivered to combustor 56. The supplemental power output required fromfirst engine system 18A between 85%-115% SHP may be beyond what ICengine 24 alone is normally capable of providing. Accordingly, thesupplemental power output may be provided by way of combustor 56.

Compared to FIG. 4 , the engine configuration represented in FIG. 5 mayprovide a lower maximum SHP but may offer better SFC when operating atLRC (˜60% SHP) with a smaller compressor 36 and smaller fuel injector(s)30A, both designed for MCR rating of 85% SHP. In other words, the engineconfiguration represented in FIG. 4 may be designed for a higher maximumSHP (i.e., +30% over the TO rating), with relatively good LRC SFC. Onthe other hand, the engine configuration represented in FIG. 5 may bedesigned for better LRC SFC, with lower maximum SHP (i.e., +15% over theTO rating).

FIG. 5 also indicates sizing requirements for various components offirst engine system 18A. As explained above the architecture of firstengine system 18A may make the supplemental power output capacityavailable without incurring significant size or weight penalty for firstengine system 18A. For example, the supplemental power capacity may beprovided separately from IC engine 24 so that the power output capacityof IC engine 24 need not be increased. Consequently, IC engine 24 andits associated systems may be designed and sized to function effectivelyand efficiently in first regime R1 and second regime R2 between 0% SHPand 85% SHP. First engine system 18A may nevertheless provide thesupplemental power output capacity using combustor 56 when required.Accordingly, as shown in FIG. 5 , some components may only need to besized for 85% SHP (i.e., MCR rating), some components may need to besized for 100% SHP (i.e., TO rating), and some components may require tobe sized for 115% SHP.

The upsizing of muffler 66 may be optional since muffler 66 couldpotentially be bypassed during the momentary operation beyond 100% SHP.Variable speed transmission 48A and gearbox 22 may be designed mainlyfor 100% SHP but may also have the capacity to momentarily operate at115% SHP with some design adjustments (if required). Similarly,components and systems associated with IC engine 24 may be designedmainly for 85% SHP but may also have the capacity to momentarily operateat 100% SHP to handle the TO phase with some design adjustments (ifrequired). In reference to FIG. 5 , combustor 56 may be used between 85and 115% SHP. Fuel Injector(s) 30A for injecting WF1 into IC engine 24may be sized/optimized to cause IC engine 24 to deliver 85% SHP. Theextra output power beyond 85% SHP may then come solely from the use offuel injector(s) 30B and combustor 56. Some components such as IC engine24, gearbox 22, variable speed transmission 48A, and optionally exhaustduct 64, and muffler 66 may still need to be sized for 100% SHP toprovide an adequate durability to permit IC engine 24 to be operatedcontinuously at 85% SHP (MCR) for example.

The embodiments described in this document provide non-limiting examplesof possible implementations of the present technology. Upon review ofthe present disclosure, a person of ordinary skill in the art willrecognize that changes may be made to the embodiments described hereinwithout departing from the scope of the present technology.Modifications could be implemented by a person of ordinary skill in theart in view of the present disclosure, which modifications would bewithin the scope of the present technology.

What is claimed is:
 1. An aircraft power plant comprising: an internalcombustion (IC) engine using intermittent combustion during operation,the IC engine being drivingly connected to a load via a gearbox totransfer motive power from the IC engine to the load, the IC enginehaving an exhaust outlet for discharging a flow of first exhaust gas; acombustor including: a combustor inlet in fluid communication with theexhaust outlet of the IC engine to receive the flow of first exhaust gasfrom the IC engine into the combustor; a fuel injector for injectingfuel into the combustor in which the fuel is mixed with the firstexhaust gas and ignited to generate a flow of second exhaust gas; and acombustor outlet for discharging the flow of second exhaust gas; aturbocharger associated with the IC engine, the turbocharger including:a turbine in fluid communication with the combustor outlet and driven bythe flow of second exhaust gas, the turbine being drivingly connected tothe load to transfer motive power from the turbine to the load; and acompressor drivingly connected to the turbine and driven by the turbineto compress combustion air for the IC engine; and a continuouslyvariable transmission, the turbine being drivingly connected to the loadvia the continuously variable transmission and the gearbox.
 2. Theaircraft power plant as defined in claim 1, comprising a shaft drivinglyconnecting the turbine and the compressor, the turbine being drivinglyconnected to the load via the shaft.
 3. The aircraft power plant asdefined in claim 1, wherein: the IC engine includes a combustionchamber; and the exhaust outlet of the IC engine is configured todischarge the flow of first exhaust gas from the combustion chamber ofthe IC engine.
 4. The aircraft power plant as defined in claim 1,wherein the IC engine is a Wankel engine.
 5. The aircraft power plant asdefined in claim 1, comprising a liquid cooling system for cooling theIC engine, the liquid cooling system including: a heat exchanger topromote heat transfer from a coolant fluid carrying heat from the ICengine to ambient air; and a fan mechanically driven by the IC enginevia a variable speed transmission to drive the ambient air through theheat exchanger.
 6. The aircraft power plant as defined in claim 1,wherein: the IC engine is a first IC engine; and the aircraft powerplant includes a second IC engine drivingly connected to the load.
 7. Ahelicopter comprising the aircraft power plant as defined in claim 6,wherein the load includes a main rotor of the helicopter.
 8. An aircraftcomprising the aircraft power plant as defined in claim
 1. 9. Theaircraft power plant as defined in claim 1, comprisingvariable-orientation guide vanes configured to guide a flow of ambientair being delivered to the compressor.
 10. The aircraft power plant asdefined in claim 1, comprising a muffler in fluid communication with theturbine to permit the flow of second exhaust gas flowing through theturbine to be discharged via the muffler.
 11. The aircraft power plantas defined in claim 1, comprising an intercooler disposed between thecompressor and the IC engine to remove heat added to the combustion airby the compressor.
 12. The aircraft power plant as defined in claim 5,comprising variable-orientation guide vanes guiding a flow of ambientair being delivered to the fan.